To date there has been no successful technique for repair of superalloy structures having holes therein which does not rely upon a static, mechanical bond joint. While the present invention was developed for specific application in the gas turbine engine field, and has particular application in this field, the invention is not limited thereto.
Gas turbine engines are widely used and well developed generators of power used primarily for propulsion of aircraft. The temperatures and stresses to which most gas turbine engine components are subjected require that such components be fabricated of high strength, high temperature materials, such as superalloys and titanium alloys. Further, the temperatures at which such components operate, and the stresses to which they are subjected, require that the materials be processed by hot working to achieve the necessary properties. Such hot worked materials have been very difficult to repair in the past without severe loss of high temperature capability. In addition, high volume fraction gamma prime nickel-base superalloys are not weldable, due to strain age cracking.
In large gas turbine engines the compressor and turbine sections are of axial flow design, and comprise a plurality of stages, each consisting of a disk or ring having a plurality of airfoils mounted on its rim. The blades and the disk or ring are mounted on rotating shafts and are subject to severe environmental conditions. Historically, blade and disk assemblies have been produced from separate components, and have been mechanically attached. While this permits manufacture of blades and disks from different materials, the use of mechanical attachment means adds substantially to the weight of the assembly.
Increases in performance requirements for gas turbine engines have now led to the development of integrally bladed rotors, wherein the blades are an integral part of the rotor, either formed integrally by such processes as isothermal forging under conditions of low strength and high ductility, or metallurgically bonded to the disk. Either form reduces the weight penalty resulting from prior art mechanical joining procedures.
With the increasing demand for higher performance and lighter weight assembly, many gas turbine components, such as the turbine section, are now produced as a one piece assembly. The primary advantages of such assemblies are in weight and fuel savings, and component life as a result of the elimination of life-limiting stress concentrators associated with mechanically attached components.
In assembly of a gas turbine engine, components are frequently joined by bolting one subassembly to another, using highly precise drilling to properly locate bolt holes in mating components. When such a hole is mislocated, misaligned, or misdrilled, or damaged during assembly or use, the subassembly in which it is located frequently must be scrapped. While this may be costly in the case of individual rotors, it is even more so in the case of integrally bladed rotors, where substantial expense has been incurred prior to joining one component to another. Thus, as the value of the individual assembly is increased, the margin of error in machining and joining components becomes more critical.
Past techniques for repair of such holes have frequently depended upon mechanical means, and have generally been limited to materials of lesser strength and temperature capability than the superalloys and titanium utilized in present day gas turbine engines. For example, U.S. Pat. No. 2,010,569, of Sitzler, teaches a method for plugging holes in plates using a special form or shape of a plug, having a cylindrical body portion and a centrally depending projection which increases in diameter from the upper end to the lower end. The plug is force fit into the hole, which has been countersunk so as to have projecting shoulders designed to lock the plug in place when pressed. The plug and metal plate are typically of the same metal, suitable for use in steam platens, and are not subjected to any heat treatment.
Crossman et al, in U.S. Pat. No. 3,952,395, teach a method for closing a drilled hole in a pneumatic or hydraulic system by press-fitting a ball into the hole to form a seal, and then "staking" the ball in position so as to crimp a portion of the work piece over the surface of the ball. In this case, the ball is of a harder material than the work piece, and is sized slightly larger than the diameter of the hole. No heat treatment is utilized.
Also, in U.S. Pat. No. 3,522,648, Weber teaches sealing a hole by deforming a ductile ball or slug of metal. Here, the spherical ball is deformed or upset so as to form a close contact and a mechanical metal-to-metal contact bond with the cylindrical walls of the hole. The ball is selected to be slightly more ductile than the casting so as to be readily deformed, and no heat treatment is used.
In U.S. Pat. No. 3,487,530, Ely teaches a method for repair of casting defects wherein one drills out the defect, fills the drilled area with a plug, diffusion bonds the casting and the plug, and removes excessive plug material. This method is specifically designed for cosmetic or surface repair of superalloy and refractory metal castings to avoid cracking problems associated with conventional fusion processes. In one embodiment of the process, the Patentee teaches boring or drilling the defect to reform the same into a truncated cone-shaped opening or hole, and driving a complementally tapered plug into the hole with a hammer to seal it therein. In all instances, a diffusion bond is induced without reaching the melting point or exceeding the yield point of the cast member or the plug. The process may also be practiced in the form of friction welding.